(1) Field of the Invention
This inventions relates to methods for applying thermal barrier coatings and, more particularly, to methods for applying a hybrid thermal barrier coating.
(2) Description of the Related Art
Gas turbine engines are well developed mechanisms for converting chemical potential energy, in the form of fuel, to thermal energy and then to mechanical energy for use in propelling aircraft, generating electrical power, pumping fluids, etc. At this time, the major available avenue for improved efficiency of gas turbine engines appears to be the use of higher operating temperatures. However, the metallic materials used in gas turbine engines components are currently very near the upper limits of their thermal stability. In the hottest portion of modern gas turbine engines, metallic materials are used at gas temperatures above their melting points. They survive because they are air cooled. But providing air cooling reduces engine efficiency.
Accordingly, there has been extensive development of thermal barrier coatings for use with cooled gas turbine aircraft hardware. By using a thermal barrier coating, the amount of cooling air required can be substantially reduced, thus providing a corresponding increase in efficiency.
Turbine blades and vanes are two exemplary cooled gas turbine aircraft components utilizing thermal barrier coatings. Turbine blades and vanes in the hot section of the cooled gas turbine are typically coated with a metallic and/or ceramic thermal barrier coating to increase their durability. Ceramic thermal barrier coatings are applied anywhere from 0.5 to 10 mils or more and can reduce temperatures at the surface of the metal substrate by up to 300° F. or more. For typical current thermal barrier coating systems, the ceramic material is applied by either a plasma-spray, typically in air (APS) process, or a physical vapor deposition process, such as electron beam physical deposition (EB-PVD).
A typical distress mode, exhibited by blades and vanes during their part life, is an oxidation prone hot spot developing on the pressure side of the airfoil. Presently, these hot spots may be coated with a low thermal conductivity thermal barrier coating, e.g., which exhibit a thermal conductivity value that is 50% to 60% lower than current commercially available thermal barrier coatings such as 7YSZ. However, as blades and vanes experience severe operating conditions at high mach numbers, another distress mode exhibited in high mach number regions of the airfoil is particulate erosion, e.g., of the coating and in cases the underlying substrate.
Consequently, there exists a need for a thermal barrier coating which provides the requisite thermal conductivity necessary to prevent oxidation yet also exhibits erosion resistant properties to withstand particulate erosion.